As is well known, the aeronautical industry requires structures which, on the one hand, support the loads to which they are submitted, meeting high stiffness and stress requirements, and on the other hand, are as light as possible. A consequence of this is the increased use of composite materials, especially CFRP (Carbon Fibre Reinforced Plastic), in primary structures due to the significant weight loss achieved compared with metallic materials.
Following this trend, there are known, for example, aircraft lifting surfaces which consist of two torsion boxes (on the right hand and left hand sides) joined to a central box manufactured entirely made out with CFRP panels, using as skins for said boxes individual pieces, that is to say, using four complete skins (two skins on top and two on the bottom) to make up the left hand and right hand torsion box. As it can be well understood, the integration of these kinds of parts presents different problems due to their great size and their complex geometry. One of these problems is the introduction of a bevelled finish on the bottom skin edge region of the lateral boxes to facilitate their union to the central box by an intermediate union plate.
The solution to this problem when the skins were manufactured with metallic materials consisted of a machining operation in said edge region to produce said bevelled finish, but this is not applicable to a composite material part.
The present invention is focused towards the solution of this problem.